The subject matter disclosed herein generally relates to gas turbine engines and, more particularly, to a method and apparatus for fuel injection control in combustors of gas turbine engines.
A gas turbine engine, typically used as a source of propulsion in aircraft, operates by drawing in ambient air, combusting that air with a fuel, and then forcing the exhaust from the combustion process out of the engine. A fan and compressor section, having a low and high pressure compressor, rotate to draw in and compress the ambient air. The compressed air is then forced into the combustor, where it is split. A portion of the air is used to cool the combustor while the rest is mixed with a fuel and combusted.
The combustor includes fuel injector, which is a device for dispersing fuel into the combustor to be combusted. The fuel enters a nozzle which atomizes the fuel to allow for greater air-fuel mixing before the combustion process. Conventionally, once the fuel injector is designed and installed, there is little ability to change performance aspects of the fuel injector, such as, for example, fuel spray angle, fuel spray atomization, and fuel spray breadth.